担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）   出版者・発行元：IOP PUBLISHING LTD  

© 2019 IOP Publishing Ltd. The results of experimental and computational studies are considered on a near surface electric discharge effect on supersonic airflow near a 15° compression surface. The tests were performed at Mach number M = 2, stagnation pressure P 0 = 1-2.8 bar, stagnation temperature T 0 = 290-600 K, and plasma power W pl = 6-12 kW. They demonstrated a significant effect of plasma on the flow structure and reduction of static pressure on the compression surface. Transient phenomena were analyzed and it was found that the pressure decrease on the ramp was as fast as t < 0.3 ms. Simulations based on 3D unsteady Navier-Stokes equations with plasma modeled as an array of lengthwise heat sources demonstrated adequacy of such simplifications. Further simulations attempted to find an optimal range of plasma power and position in terms of achievable effect, effectiveness of the method, and response time of the system to the plasma actuation. The electric discharge authority for a fast and effective control of aerodynamic forces in a compression ramp configuration is considered.

DOI： 10.1088/1361-6463/ab352f

記述言語：英語   掲載種別：研究論文（学術雑誌）   出版者・発行元：MDPI Multidisciplinary Digital Publishing Institute  

This study considers the effect of an electric discharge on the flow structure near a 19.4° compression ramp in Mach-2 supersonic flow. The experiments were conducted in the supersonic wind tunnel SBR-50 at the University of Notre Dame. The stagnation temperature and pressure were varied in a range of 294-600 K and 1-3 bar, respectively, to attain various Reynolds numbers ranging from 5.3 × 10 5 to 3.4 × 10 6 based on the distance between the exit of the Mach-2 nozzle and the leading edge of the ramp. Surface pressure measurements, schlieren visualization, discharge voltage and current measurements, and plasma imaging with a high-speed camera were used to evaluate the plasma control authority on the ramp pressure distribution. The plasma being generated in front of the compression ramp shifted the shock position from the ramp corner to the electrode location, forming a flow separation zone ahead of the ramp. It was found that the pressure on the compression surface reduced almost linearly with the plasma power. The ratio of pressure change to flow stagnation pressure was also an increasing function of the ratio of plasma power to enthalpy flux, indicating that the task-related plasma control effectiveness ranged from 17.5 to 25.

DOI： 10.3390/aerospace6030035

掲載種別：研究論文（国際会議プロシーディングス）  

© 2018 IEEE. Demand for aircraft transportation has doubled in the past ten years and is expected to increase. On the other hand, electrification of the aircraft's equipment is gradually improving. As part of this electrification, mounting of electrically driven wheel is considered. Our previous research also proposed safety landing by electrically driven wheel. In this paper, we propose the total thrust control method using propeller and electrically driven wheel in ground run. This method can accelerate in constant value on dry or wet road surface. We demonstrate the effectiveness of the proposed method by simulations and basic experiments.

DOI： 10.1109/AMC.2019.8371103

記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）   出版者・発行元：IOP PUBLISHING LTD  

© 2017 Published under licence by IOP Publishing Ltd. An experimental research has been conducted to investigate the flow control method using opposing jet with air and water as a fluid spike on a bunt body in hypersonic freestream at Mach number 7, using the Kashiwa hypersonic wind tunnel. The dynamic characteristics and instability of opposing jet were clarified. It can be seen from the comparison between the air opposing jet and the water opposing jet that a liquid jet seems more effective in generating fluid spike for reduction in the drag force and the heating rate on the body. Details of the flow field and the jet oscillation have been captured by Schlieren system with high-speed cameras. The experimental data reveal the interaction of the jet oscillation with the shock wave fluctuation. Such interaction has strong impact on the whole flow field structure and should be taken into consideration in the design of the active flow control system using the opposing jet method.

DOI： 10.1088/1757-899X/249/1/012014

担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）   出版者・発行元：IOP PUBLISHING LTD  

© 2017 Published under licence by IOP Publishing Ltd. This study examined the effect of Reynolds number on plasma-assisted flow control ahead of a compression ramp geometry in Mach-2 supersonic flow. The experiments were conducted in the supersonic wind tunnel SBR-50 at the University of Notre Dame. Stagnation temperature and pressure were varied as T0=294-500K and P0=1-3bar to attain Reynolds number ranging from 3.4×105-2.2×106. Ramp pressure measurements, schlieren visualization, and high-speed camera imaging were used for the evaluation of plasma-assisted flow control effects. A linear dependency was found between the ramp pressure change per averaged plasma power and Reynolds number.

DOI： 10.1088/1757-899X/249/1/012006

記述言語：英語   掲載種別：研究論文（国際会議プロシーディングス）   出版者・発行元：IOP PUBLISHING LTD  

© 2017 Published under licence by IOP Publishing Ltd. Considering a spacecraft that encounters particle-laden environment, such as dust particles flying up over the regolith by the jet of the landing thruster, high-speed flight of a projectile in such environment was experimentally simulated by using the ballistic range. At high-speed collision of particles on the projectile surface, they may be reflected with cracking into smaller pieces. On the other hand, the projectile surface will be damaged by the collision. To obtain the fundamental characteristics of such complicated phenomena, a projectile was launched at the velocity up to 400 m/s and the collective behaviour of particles around projectile was observed by the high-speed camera. To eliminate the effect of the gas-particle interaction and to focus on only the effect of the interaction between the particles and the projectile's surface, the test chamber pressure was evacuated down to 30 Pa. The particles about 400μm diameter were scattered and formed a sheet of particles in the test chamber by using two-dimensional funnel with a narrow slit. The projectile was launched into the particle sheet in the tangential direction, and the high-speed camera captured both projectile and particle motions. From the movie, the interaction between the projectile and particle sheet was clarified.

DOI： 10.1088/1757-899X/249/1/012015

担当区分：筆頭著者   記述言語：英語   掲載種別：研究論文（学術雑誌）  

© IMechE 2016. Breathing blunt nose technique is one of the promising methods for reducing the drag of blunt-nosed body at hypersonic speeds. The air, traversed by the bow shock positioned ahead of the nose, at the stagnation region is allowed to enter through a hole at the blunt-nose and ejected at the rear part (base region) of the body. This manipulation reduces the positive pressure over the stagnation regions of the nose and increases the pressure at the base, resulting in reduced suction at the base. The simultaneous manifestation of reducing the compression at the nose and suction at the base regions results in reduction of the total drag. The drag reduction caused by the breathing blunt nose technique has been measured in a Mach 7 tunnel. Also, the drag and flow field around the blunt-nosed body, with and without breathing hole, has been computed. The aerodynamic characteristics of the breathing blunt nose model obtained experimentally are compared with the CFD results. It is found that the breathing results in 5% reduction in drag. The lift coefficient also comes down for the model with breathing nose. But the lift-to-drag ratio is found to be the same for both the cases; the blunt-nosed body with and without nose-hole.

DOI： 10.1177/0954410016643979

掲載種別：研究論文（学術雑誌）  

© 2016 Author(s). In this paper, we present a novel method for analyzing electrohydrodynamic (EHD) thrusters. The method is based on a perturbation technique applied to a set of drift-diffusion equations, similar to the one introduced in our previous study on estimating breakdown voltage. The thrust-to-current ratio is generalized to represent the performance of EHD thrusters. We have compared the thrust-to-current ratio obtained theoretically with that obtained from the proposed method under atmospheric air conditions, and we have obtained good quantitative agreement. Also, we have conducted a numerical simulation in more complex thruster geometries, such as the dual-stage thruster developed by Masuyama and Barrett [Proc. R. Soc. A 469, 20120623 (2013)]. We quantitatively clarify the fact that if the magnitude of a third electrode voltage is low, the effective gap distance shortens, whereas if the magnitude of the third electrode voltage is sufficiently high, the effective gap distance lengthens.

DOI： 10.1063/1.4951721

担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）   出版者・発行元：一般社団法人 日本航空宇宙学会  

As a part of the "Membrane Aeroshell for Atmospheric-entry Capsule" (MAAC) project pursued by the University of Tokyo, JAXA and several other universities, a demonstration is planned to deploy an atmospheric-entry nano-satellite from International Space Station (ISS). The effect of aeroshell deployment timing on the orbital decay was evaluated and it was clarified that the orbital decay is significantly accelerated after aeroshell deployment due to the change in its ballistic coefficient. In order to keep the forward surface of the aeroshell in the direction of motion at the moment of aeroshell deployment, it is necessary to obtain the attitude of satellite with reference to the orbital direction. The applicability of the Faraday cup as an ion detector is discussed through experimental results at Inductively Coupled Plasma (ICP) wind tunnel and it was revealed to be promising for the estimation of the satellite direction.<b> </b>

DOI： 10.2322/tastj.14.Pf_1

記述言語：英語   掲載種別：研究論文（学術雑誌）   出版者・発行元：一般社団法人 日本航空宇宙学会  

<p>To realize the penetrator mission for icy objects in the solar system, the penetration dynamics into ice has been experimentally and numerically studied. In the experiments, the ogive/cylinder projectiles of the mass about 2.6 g were tested for the target made from water ice by the ballistic range at the impact velocity 100-350 m/s. The behavior of the projectile and target was observed by the high-speed video camera. The penetration trajectory was visualized by pouring the plaster into the crater. The empirical analysis model to describe the force acting on the body surface has been developed, based on the panel method and the shock wave analogy. It includes two parameters representing the effective speed of sound of crushed ice and the damping of the pitching motion. Fairly good agreement with the experiments was obtained with respect to the stop conditions of the penetrator by setting these parameters appropriately. The computational result assuming a flight model of the mass 14.9 kg and the impact velocity 300 m/s shows that the maximum deceleration G is in the same order as that for the lunar penetrator.</p>

DOI： 10.2322/tastj.14.Pk_65

記述言語：英語   掲載種別：研究論文（学術雑誌）   出版者・発行元：IOP PUBLISHING LTD  

� 2015 IOP Publishing Ltd. This study presents a new method for predicting a precursor of breakdown voltage. This method applies a global linear stability analysis to a drift-diffusion model coupled with the Poisson equation. The instability of the solution of these equations is evaluated with the proposed method and is considered to be the onset of a breakdown. The proposed method is validated in one dimension by using the Townsend theory as a reference. To extend the one-dimensional case and investigate the characteristics of multidimensional eigenmodes, we apply the method to two-dimensional plane-to-plane discharge tubes. To prove an adaptability against complex electrodes geometry, the discharge path and the corresponding breakdown voltage for the well-known corner-to-plane geometry are qualitatively evaluated.

DOI： 10.1088/0963-0252/24/5/055014

掲載種別：研究論文（国際会議プロシーディングス）  

© 2015 IEEE. Aircrafts are desired to be more energy efficient and safer due to the increasing demand for air transportations. However, generally speaking, nowadays commercial airplanes tend to loss stability under wind disturbances, especially during landing. On the other hand, electric airplanes (EAs) are believed to satisfy the two demands because electric motors are used for the propulsion. Especially, as the actuators, electric motors can improve the control performances of EAs compared with internal combustion engines (ICEs). In this paper, by utilizing electric motors' advantages, lift control method using propeller slipstream is proposed for safe landing, which might be a key technology to innovate the design of EAs. Moreover, simulations and experiments are conducted to verify the effectiveness of the proposed method.

DOI： 10.1109/ICMECH.2015.7083998

担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）   出版者・発行元：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

In this study, the application of direct current arc plasma discharge is discussed through computational analysis on the flow around a bump-up shape over a surface of a typical hypersonic vehicle under 20-km altitude and Mach-5 flight condition. Numerical simulation based on our simple plasma model suggested that the front pressure over a bump-up slope could be decreased by the use of plasma actuation method proposed in this study. The plasma source generated at 7 cm ahead of the bump successfully reduced the drag coefficient for more than 56%, changing the surface flow to go smoothly along the surface of the bump shape.

DOI： 10.2322/tastj.12.Pe_1

記述言語：英語   掲載種別：研究論文（学術雑誌）   出版者・発行元：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

The EHD (electro-hydrodynamic) thruster seems suitable for a device of mobility at planetary exploration because of its propellant-less nature. In this paper, the fundamental characteristics of the EHD thruster are investigated from a viewpoint of the applicability to the planetary exploration. The numerical analysis of the drift-diffusion equations shows that it works irrespective to the polarity of the electrodes because the polarity of the produced ions is also changed. This fact implies that it will work in different atmospheric composition with different polarity of the produced ions. It is also numerically found that the thrust tends to increase with the decrease in the ambient pressure. For simplicity of discussion, a simple analytical model to describe the energy conversion efficiency is derived, and it indicates that higher efficiency is expected in higher Knudsen number regime. Both the above results suggest that the EHD thruster is promising especially in planetary atmosphere with low density.

DOI： 10.2322/tastj.12.Pe_5

担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）  

DOI： 10.2322/tastj.12.To_2_1

担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）   出版者・発行元：WALTER DE GRUYTER & CO  

An experimental visualization has been carried out to study the dependence of the pressure hill height and the influence zone expanse for flow past rectangular blocks of flat square face and varying length, over a Reynolds number range from 1364 to 4931. It is found that, the pressure hill length and the influence zone expanse decrease with the length to width ratio of the block, up to about L/W 1, for Reynolds number up to 1586. For higher Reynolds numbers, both H/W and Z/W increase with the model length, till L/W 1. For L/W more than 1, both H/W and Z/W gradually become independent of L/W. The ratio of Z/H is influenced only marginally by L/W up to 1, and for greater values of L/W, Z/H is almost a constant at all Reynolds numbers of the present study. Copyright © 2011 De Gruyter.

DOI： 10.1515/TJJ.2011.067

担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）  

DOI： 10.1615/VisMechProc.v1.i3.50

担当区分：筆頭著者,　責任著者   記述言語：英語   掲載種別：研究論文（学術雑誌）   出版者・発行元：THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES  

A new idea of an aerodynamic control device for hypersonic vehicles using plasma discharges is presented. The effect of DC plasma discharge on a hypersonic flow is examined with both experiments and CFD analyses. It is revealed that the surface pressure upstream of plasma area significantly increases, which would be preferable in realizing a new aerodynamic control devices. Such pressure rise is also observed in the result of analyses of the Navier-Stokes equations with energy addition that simulates the Joule heating of a plasma discharge. It is revealed that the pressure rise due to the existence of the plasma discharge can be qualitatively explained as an effect of Joule heating.

DOI： 10.2322/tastj.8.Pe_15

記述言語：日本語   掲載種別：記事・総説・解説・論説等（学術雑誌）   出版者・発行元：一般社団法人 日本航空宇宙学会  

東京大学柏キャンパスに設置されている極超音速高エンタルピー風洞について紹介する.1つの空気加熱器に対し,極超音速風洞と燃焼風洞の測定部を有する1加熱器2運転モードを採用していることが最大の特徴である.設備概要と性能,気流特性に加え,1)広範囲での融合的研究の場とするため実験は公募制とし,外部ユーザーも積極的に受け入れること,2)大学設備としての安全性と低コスト性,高運転頻度を実現すること,3)大学の研究室として持続的に管理ができること,を旗印にした運用についても説明する.設備情報や行われてきた実験プロジェクト内容はホームページで公開されており,外部ユーザーの便宜が図られている.さらに,極超音速・高温気流実験装置の利用範囲を広げていくためのきっかけとなるものと期待して,分野融合的な実験例も紹介する.

DOI： 10.14822/kjsass.63.7_223

記述言語：日本語   出版者・発行元：日本航空宇宙学会  

記述言語：日本語   出版者・発行元：日本航空宇宙学会  

記述言語：日本語   出版者・発行元：日本航空宇宙学会  

© 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All Rights Reserved. Flow control method using direct current (DC) plasma discharge in Mach-7 hypersonic flow was investigated as a method to improve aerodynamic characteristics of hypersonic vehicles. In this work, the effect of DC arc plasma discharge was investigated to clarify its flow control effectiveness both with numerical analysis and with hypersonic wind tunnel experiments. When a plasma discharge is maintained ahead of double-wedge shaped surface that represents a body flap of hypersonic vehicles, numerical simulation based on simple energy addition model that represent plasma was conducted and it revealed that the aerodynamic force acting on the slope significantly decreases by more than 10 percent. Experimental results also suggested that the force acting on the slope decreases a lot within a very short time less than 0.1 seconds, suggesting that the present plasma discharge configuration can serve as an efficient aerodynamic control technique with a very short response time.

DOI： 10.2514/6.2015-2342

© 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. In this study, the formation of bow-shock and its instabilities have been studied in front of axisymmetric blunt nose shapes, at hypersonic Mach number 7 to understand the physics of bow-shock instabilities. The three geometries have been investigated as concave hemispherical shell (with negative curvature), convex hemispherical shell (with positive curvature) and circular at plate (with zero curvature), with same characteristic diameter. The lift and drag force measurements and unsteady pressure measurements were performed for all three geometries. The bow-shock visualizations have been carried out by Schlieren system, using high-speed camera at 10000 fps in hypersonic wind tunnel. By using image processing method, bow-shock unsteadiness has been analyzed for all three geometries. The large deformations in bow-shock in front of hemispherical concave shell have been analyzed with unsteady pressure measurement and shock displacements at the center and the edges of the geometries. It is observed that the large bow-shock deformations for convex, at and concave shaped blunt nose sustain within time order of 0.1 ms, 0.5 ms and 100 ms to 200 ms, respectively.

記述言語：日本語   出版者・発行元：日本航空宇宙学会  

Hypersonic flow control with plasma discharge has received a good deal of research attention in recent years. In this study, the direct current (DC) plasma discharge over a flat plate was investigated as a prospective futuristic aerodynamic control device. Spectroscopic analysis showed that the vibrational temperature of nitrogen molecule is 7000 K and the translational temperature is 3500 K at maximum, which suggested that the plasma is at a vibrationally nonequilibrium state. Since electromagnetic forces for the futuristic aerodynamic control can mainly be imposed on ionized molecules, a numerical analysis was conducted based on Park's Two-temperature model and Gupta's 11-species model in order to estimate the distribution of high vibrational-temperature region. It was revealed that the concentration of high-vibrational temperature region exists in the boundary layer over flatplate surface, which is consistent with the experimental result of vibrationally exited nitrogen molecular distribution. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Flow control with plasma discharge is one of the promising techniques for the futuristic aerodynamic control methods. In the present work, a hypersonic flow control method which utilizes a direct-current (DC) arc plasma discharge is investigated as a fundamental study on such a new device. Hypersonic wind tunnel experiments were conducted to measure the effect of the discharge on Mach-7 flow over a compression and a expansion corner. The experimental results showed that, in the case of the expansion corner, a large pressure fall induced by the plasma discharge was observed. However, the change in the entire flow field such as shock wave formation was not clear. In the case of the compression corner, the boundary separation at the corner is enhanced and enlarged by the discharge, causing the pressure to change near the plasma area. Numerical analyses were conducted on both of these cases to understand the mechanisms of the observed phenomena. Park's two temperature model and Gupta's 11-species model were employed to describe the plasma flow. In order to simulate the effects of the plasma, a simple energy addition model is proposed, where the discharge effect is replaced by an energy addition to both the translational-rotational and the vibrational-electron-excitation energy modes. The numerical results based on the proposed model successfully showed a good qualitative agreement with the experimental results for the nitrogen molecular light emission. It was revealed that the present idea of utilizing the DC arc discharge as a flow control method is more effective in manipulating the flow when it is applied to a compression corner.

DOI： 10.2514/6.2013-3130

Chemical reactions in hypersonic flows have a potential to produce much more various species than those in a static chamber because of its nonequilibrium nature. In the present study, we propose a new experimental technique of a "chemical factory" in a hypersonic wind tunnel flow. The hypersonic chemical factory is composed of a model with ablative part as a materials supplier, a pair of electrodes for electric discharge as an energy supplier, and a wake flow behind the body in frozen chemistry as a products transporter. To demonstrate this concept and to experimentally simulate the formation of prebiotic materials in high temperature shock layer around an ablating icy object entering the early earth's CO2/N2 atmosphere, a model with an ablative nose part made from water ice and dry ice is tested in the hypersonic wind tunnel at Mach number 7 in Kashiwa campus, the University of Tokyo. When the plasma discharge is ignited between the electrodes located on the side surface of the model, we observed the emission probably from CN species produced from the N element in the air and the C element in the ablation gas. The presence of CN suggests the production of HCN, which is one of the most important prebiotic materials, in hypersonic flows around extraterrestrial icy entry objects on the early earth. On the other hand, no signal of CN is observed in case of the nose made from water ice only. To assist the design of the experimental setup and to deepen our understanding on the phenomena, the thermochemical nonequilibrium Navier-Stokes analysis with C-H-O-N 28 species is also conducted. The results show reasonable agreement with the experimental observation. The present technique of the hypersonic chemical factory seems promising for investigation on nonequilibrium chemical process in a flow.

担当区分：筆頭著者,　責任著者  

The application of plasma discharge to hypersonic flow control methods has received a good deal of research attention in recent years. As a fundamental study for such methods, direct current plasma discharge in Mach-7 hypersonic flow over flat plate was investigated both with wind tunnel experiments and numerical analyses. Impulsive discharge in a boundary layer over flat plate caused a decrease in the wall pressure around electrodes. For the purpose of clarifying the principal cause of this phenomenon, a numerical analysis was conducted by solving the three-dimensional Navier-Stokes equations with impulsive energy addition to the flow. Numerical results revealed that the pressure oscillates in the vicinity of heated region, which suggests that the fluctuation of surface pressure observed in the experiment could be qualitatively explained as the unsteady response of the flow to the impulsive Joule heating by the discharge. © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

DOI： 10.2514/6.2011-3736

As a preliminary study on the application of plasma discharge to hypersonic aerodynamic control, the effect of direct current (DC) plasma in a hypersonic flow was examined with wind tunnel experiments. It was shown that the pressure rises significantly in the upstream region of the discharge. To the contrary, the pressure decreases in the region very close to the electrodes. This pressure change would be applied to a new aerodynamic control device. Under the assumption that the effects of the plasma discharge in a hypersonic flow are mainly due to the Joule heating, three-dimensional Navier-Stokes analyses were conducted. It was concluded that the pressure change far from the plasma area can be explained as an effect of the heat addition to the flow. © 2010 by the American Institute of Aeronautics and Astronautics, Inc.

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